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What would a Category II ILS approach require that a Category III approach wouldn't?
According to @RalphJ's first comment on my earlier question about ILS approaches in an E-Jet sans pitch trim:

Are you sure that Cat III approaches are allowed in that condition? Typically, Cat III requires basically everything to be operative, unless there is some pretty unusual redundancy in the aircraft. Wouldn't be surprised if one channel out removes Cat III, and then losing the 2nd channel takes out Cat II. Just a guess, but the list of things required for Cat II that aren't needed for Cat III, while not empty, isn't very long in most cases. [Emphasis (both bolding and italics) in first sentence present in original; emphasis in second and last sentences mine.]

"Requires basically everything to be operative" seems like basically the highest bar possible in terms of functional functionality, so that's kinda difficult to square with it being possible for something to not be required for an ILS-III approach despite being required for an ILS-II approach.
What sort of things would an aircraft require to be operative for an ILS-II approach, yet not need in working order for the more-demanding ILS-III approach?

If approach lights (ALSF-II) go inop when inside the FAF, it's a required go-around Cat II, because at 100' you're looking at them + the runway. Same failure during a Cat III, you can continue, because at a 50' DA the approach lights are all behind you & only the runway lights matter.
Not all OpSpecs may offer that relief, but I've seen one that does. AFAIR, that's what I had in mind with that statement in the other thread.

Do you need even the runway lights if you're going Cat IIIb/c, where there's no DH at all and you ride the ILS all the way to touchdown and beyond?

IIIB you still have to steer off the runway, so probably required there. IIIC, no idea what those OpSpecs look like.  In really, really low vis, the taxiing can be as challenging as the approach, and with less automation to help out.


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Formula for comparing drag on planes flying at different speeds at different altitudes. Ex: Mach .80 at 40,000 ft, vs Mach 5 at 100,000 ft
I am interested in getting a sense of how much extra energy is needed to overcome drag at different speeds and altitudes. For example, for me a baseline would the drag on an airplane flying something like Mach .80 at around 40,000 ft. Then I would be interested to understand, that same exact airplane, how much energy would be required if it was more like a Concorde and flying at Mach 2 at 55,000 ft, or when we read about these sci fi hypersonic transports that Boeing and Airbus theoretically research, how much energy is required to overcome something like Mach 5 at 100,000 ft?
I found a table that gives various measures for densities and pressures at various altitudes (https://www.engineeringtoolbox.com/standard-atmosphere-d_604.html)
I can express the speeds I am interested in as Mach, knots, km/h, etc.
Is there a simplified formula to be able to make these sorts of calculations?

No, especially not for the conditions you stated.
Even if we consider that "same exact ariplane" will not melt away at Mach 2 or 5, different phenomena start to appear at such speeds.
Let's restrict speeds to below about Mach 0.5 for now. At such speeds, compressibility of air does not affect characteristics significantly, and things become easier. Still, drag will depend on how exactly the airplane is built.
Most aerodynamic forces at such speeds, including drag, are proportional to dynamic pressure:
$$q = \frac{\rho V^2}{2}$$
That is, linearly proportional to air density and to square of speed.
Now, drag consists of two parts: one related to lift, called induced drag, and everything else, often called parasitic drag. This parasitic drag, at such low speeds, will indeed change proportional to $q$,1 and can be calculated easily using the standard atmosphere table you found.
But not so with induced drag. Your lift will also change with $q$, and if you want to fly level, you don't want that. So you aim to have the same lift. You do it by adjusting the angle of attack. But by doing that, you change the "aerodynamic efficiency" of the airplane: that is, you change induced drag disproportionally. Normally, there is one "best" angle of attack, and if you move away from it, you'll have more induced drag even for the same lift. How much more? This entirely depends on how the wing and the entire airplane is designed.
You can find conditions where $q$ will be the same: you can fly higher (=> lower density) and faster, and come up with the same dynamic pressure. Pilots will see it as the same indicated airspeed (IAS). In such conditions, you will fly at the same angle of attack, and drag will be the same.2
So, in general, some of the drag will change predictably in a more or less universal manner, and some will depend on the airplane. And this is for the trivial case of low speeds.
Now enter compressibility (and other) effects, and things get much more complicated. Even the parasitic drag will change non-linearly; it will acquire a component called wave drag, which depends not on $q$ but on the Mach number - and on the actual shape of the airplane. Some airplanes will not see much up to M0.8 or even more, some will suffer at 0.6. The effect is most complicated at transonic speeds; at high supersonic speeds, shape starts to matter less. Furthermore, the way lift is generated also changes, which may produce extra drag (see e.g. trim drag).
So, your best bet is to get approximate data for some "typical" airplane. Concorde is a good candidate: it has some published data regarding its efficiency at sub- and supersonic speeds.

1 To the first level of approximation: there are still some fundamental effects that change, but we are doing back-of-an-envelope calculations.
2 Note though that on a practical level it will still affect the "extra energy needed". First, this energy comes from an engine, and its efficiency changes not only with $q$. Normally, it becomes more efficient in colder air (i.e. higher). Second, you'll reach the destination faster when flying higher at the same $q$ (IAS) and thus at about the same power setting (yet, on the other hand, having spent some energy climbing).

Thanks great answer. I am still left hungry to understand exactly how much more energy per seat mile is required for a typical subsonic airliner vs. Concorde/Boom Overture vs. Boeing/Airbus Mach 5+ hypersonic concepts. I know from experience with more common existing airplanes that there are circumstances where you are going to fly your CRJ900 at 450 kn at lower seat-mile fuel flow than your Q400 at 315 kn. So going higher is definitely helpful.

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How would a new "clean sheet" General Aviation piston engine with all the latest technology differ from existing Continental/Lycoming types?
My understanding is that current GA piston engines from Continental and Lycoming are based on a design from the 1960's. There have been many evolutionary improvements, but only within the limitations of those 1960's designs. They are large displacement (~1.5L/cylinder), run at RPMs in the mid 2000s range so that they can spin the prop at optimal rates without using a gearbox, and they are air cooled to save the weight and complexity of a liquid cooling system.
What would a completely new, designed from the ground up GA piston engine look like today? Factoring the massive improvements that have been made in all sorts of engines over the years in non-aviation applications- for autos, trains, boats, generators, etc.
Would they still operate at such low RPMs? Yes reliability and durability are still major concerns, but manufacturing tolerances are so many times smaller now then they were in the 1960s (possibly allowing for reliable operation at higher RPMs), and they are now mounting fixed gearboxes to commercial airliner turbine engines so adding a fixed gearbox to a GA piston engine should be trivial.
Would they still be air cooled? Yes it can still save weight, but possibly at the cost of extra cooling drag, and definitely at the cost of much more finicky operation that requires the pilot to carefully monitor engine temps throughout their flight cycle.
Would they still be in a flat engine arrangement (flat 4, flat 6, etc.)? I would imagine so, I can't imagine an alternate piston engine arrangement with the smoothness, packaging, and scalability advantages of the flat architecture. Straight engines get too long, and V engines require many more changes to scale the number of pistons up or down.
Would they still fly on spark-ignited gasoline-type fuel, or could compression ignition based on jet fuel produce the power density and reliability required? I realize that the single largest factor in this area might be the simple availability (or lack thereof) of fuels, but I am talking about purely from an engineering perspective. Suppose we are talking about North America where both Avgas and Jet A are widely available, what would be the superior fuel for flying GA piston engines?
What other differences might there be that I haven't considered?

The low RPM isn't so much a function of the engine as it is the propeller. Propeller tips must be kept sub-sonic. There are clean-sheet designs in the new diesel/JET-A pistons being put in some aircraft.

@RonBeyer Yes I addressed the prop-speed aspect of the equation, however could they not merely add a fix gear to allow the engine to run at much higher RPM? According to basic engine design principles, higher RPM should allow higher power density... of course many potential downsides, some of which I mention...

I'm most certainly not an aviation engine expert, but it seems to me that running the engine at a higher RPM then gearing the output back down to what is required for the prop is simply a recipe for burning more fuel to cover the same distance - not an appealing prospect. My car has 5 gears - at 2000 RPM in 3rd I can do 55MPH, in 4th, 65MPH and in 5th, 75MPH. I can change gearing to travel faster at the same fuel/hr, but that's not how it works in a prop aircraft. If anything, you'd want the engine spinning slower than the prop to save fuel (other considerations aside).

@FreeMan: The difference is that at highway speeds, your car engine is producing only a small fraction of its rated power (all that's needed to overcome air friction & rolling resistance).  Most of the potential power is just used for acceleration.  (Which is why hybrids get better mpg.)  An aircraft engine runs at a large fraction of max power all the time.

A clean sheet design is the Jet Engine, no? Tbh, if you’re a 100lbs overweight, then the real cost savings is right under your nose.

It would have a gearbox between the crankshaft and the propeller. That is certain.
Other very likely details would be:

overhead camshafts
four valves per cylinder
water cooling at least for the cylinder heads
electronic ignition
single lever control

This would enable it to run constantly at about 4000 RPM and create the same power from a smaller displacement. The consequence would be a more compact and lighter core engine. Compared to a car engine, the cooling system would be more capable, so the weight would still be higher than that of a car engine of the same nominal shaft horsepower. But it should be lighter and allow for a more aerodynamic cowling than the current piston aero engines.
The Porsche flight motor (PFM 3200) was like that, and for a reason. Yes, it did not outperform the older engines in terms of mass, but it was noticeably less noisy. Starting from a dated six-cylinder boxer design, the engineers could not fully optimize the engine for its intended purpose.
Once a gearbox is fitted, it makes sense to reduce propeller speed, so the diameter  of the propeller fitted to that engine would be a bit larger than that of a comparable propeller on a conventional engine.
Once Jet-A is widely available at GA airfields, a diesel version would also be very likely. Note, however, that the system weight of a piston engine (engine + accessories + fuel system + fuel) will be higher for a diesel on short ranges.

Thanks for these details. The changes you mention- overhead camshafts, four valves per cylinder, electronic ignition, single lever control- are these things that are impossible to integrate as evolutionary changes to the existing Continental/Lycoming designs? Regarding 4000rpm, is that just the current limit for reliable long-term operation?

@Charles847: Overhead camshafts and four valves per cylinder were already used in WW II aero engines, so they should be no problem to integrate. The only barrier is missing competition. Electronic ignition and single-lever control would be a no-brainer for every decent engine designer as well - same problem. The incumbents are protected by a massive wall of regulation which makes them lazy. The 4000 RPM are a conservative compromise between engine life and performance, but no fixed limit.

Why Jet-A rather than plain old automotive diesel?

@jamesqf: Jet-A has a specified quality. Please read this answer for details. I could not explain it half as well.

@PeterKämpf From what I've seen around this place, if you say you couldn't explain it better, it must be a very good answer.

@Peter Kämpf: That link doesn't really explain why you'd run Jet-A in a GA diesel.  To quote from the link itself, "theres no reason to burn a more tightly specified (and thus more expensive) fuel when a less expensive one is readily available",   Plain old diesel obviously works quite well in automotive diesels: why would it not work just as well in a (properly designed) GA diesel engine?  The issues of octane ratings that apply to some older gasoline engines wouldn't apply to a newly-designed diesel.

This answer needs to explain why for each of the points. Water freezes, expands, and moves around when flying. Valves are for breathing purposes and really shine for high rpm, even still twice the valves doesn’t mean twice the air flow. OHC I don’t understand why that’s important either. Why even have a cam in the first place and just go straight electric or pneumatic?

@Mr.DemetriusMichael ok some detailing. Freezing water in the cooling system shouldnt be a problem during opreration, when the cooling water is > 90°C usually. like with cars, that might be at most a problem when parking in really cold environments ( < -20°C). OHC for sharper valve timings while using mature tech - beyond that, yeah why not electric or pneumatic. Thing is - tech maturity is often measured in thousands of kilo-operating-hours (kOh)  of field experience. I don't know where we are with electric or pneumatic (F1 .. a few hours ..)

It would look almost identical mainly because the engines are designed around a use case not for purely engineering pursuits. You may see some different material choices, some changes in piping and head design as our understanding of these topics has improved and our production abilities have come a long way but ultimately little change.
You should read the answers here, and here. They are almost duplicates but more focus more on why modern car engines are not used in aviation. Although in some cases they are.
Piston GA engines with reductive gear boxes were attempted in the 80's on the Mooney M20 PFM ultimately the project was such a failure Mooney paid to replace all the engines with continentals.
As @RonBeyer mentions in the comments the low RPM is a function of the propellor not the engine.
Aircraft engines simply have a different use case than automotive engines which has kept them simpler over the years. The cost of certification (a huge factor in aviation) has also kept up rapid innovation at bay.

or could compression ignition based on jet fuel produce the power
  density and reliability required

Some modern piston diesels are already certified to run on Jet A.
This article says it pretty much as well as I can and is worth the read.
Something that is perhaps overlooked is that aircraft engines are not subject to the same regulations that automotive engines are. As such they are, in some cases actually more efficient. One major thing that impacts car engines efficiency is catalytic converters and mufflers, something GA planes dont really have. As such there has never been a reason to compensate for the performance loss as a result of regulations.
You also mention "more reliable", GA engines are quite reliable and the FAA/NTSB has reported that mechanically related accidents are on a general down trend if we look back at history. Modern machining and industrial practices can be used to make old engines in a more reliable and consistent way.
One major efficiency boost that has been realized on the ground is turbo charging. While there are plenty of turbo aircraft out there the prevailing attitude in aviation has always been to keep things simple. Less parts means less to fail as such there is simply a smaller market for a turbo aircraft engine. It is not that the technology is a mystery to the makers or the consumers its simply an unwanted complexity.

I just finished reading a fascinating article on the engine used in your mentioned Mooney M20 PFM, and I recognize all those downsides, but they seem to be downsides inherent to the application of automotive-optimized engines to aviation, rather than downsides inherent to the application of modern piston engine technology to purpose-built aviation engines. Shouldn't advancements in our technology level allow for aviation engines that are smaller, lighter, more reliable, more efficient, for the same power, just as our tech advancements have enabled such gains for auto/marine/stationary engines?

The answers linked mention all the non-technological reasons for the engine differences, but clearly imply new design would at least use higher RPM. It also probably would use all the computer control, which is now well tested in automobile operation.

This is not the answer to the question. it's not about use of an automotive engine in GA. It's about comparing the progress that has happened in the automotive sector (which is huge), compared to the one that has happened in the GA sector.

Claiming they would look identical is like claiming that modern car engines look identical to car engines from the 60ies, just because already then they were designed for a car use case.

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It would look like the modern GA engine already in production, the Austro AE300:

a car engine basis, in the AE300 case a Mercedes diesel.
water cooling.
a 1:1.69 reduction gearbox.
running on diesel or kerosene (Jet A1).

Not sure about that. My (strong) assumption is that they use the AE300 as base because then they can save half the engineering development and testing by relying on proven parts, as a trade off to keep development costs down, especially in a small-volume market, where development costs play a larger role than in mass production market like the automotive industry.
My assumption is - if you had a sponsor who's happy to pay anything - you'd get a different design

Rotax makes modern GA engines even for certified application. But they don't differ that much as car engines from aviation engines as they work and different requirements. https://www.flyrotax.com/produkte/detail/rotax-915-is-isc.html

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Does the quarter-chord-angle rule still hold for wings with very long chords or large differences between leading- and trailing-edge sweep angles?
For most fixed-wing aircraft, the effective wing-sweep angle (i.e., the angle that the aircraft aerodynamically behaves like its wings are swept at) is equal to the angle of the quarter-chord line (the line along the wing formed by the points 25% of the way from the leading edge to the trailing edge).1  This is because the aerodynamic effects of wing sweep are a product of the spanwise flow along the wing (the airflow from root to tip, or vice versa), which is a product of both the leading-edge sweep angle and the trailing-edge sweep angle; as the contribution from the leading-edge sweep angle is somewhat greater than that from the trailing-edge sweep angle, the effective sweep angle is closer to that of the leading edge than to that of the trailing edge.
One can, however, imagine aircraft where the effects of the leading-edge and trailing-edge sweep angles on the airflow over the wing are less strongly coupled.  For an aircraft with a very long wing chord (a very large distance between the leading and trailing edges), one might reasonably expect that air passing over the wing would be affected by the leading edge, then start to return to straight-rearwards flow, and only then be affected by the trailing edge, with the great distance between the leading and trailing edges serving to effectively decouple the effects on the airflow over the wing of the two edges.  Likewise, for an aircraft with a very large difference between leading-and-trailing-edge sweep angles (imagine, for instance, an aircraft with its wings' leading edges swept rearwards at 45° and their trailing edges swept forwards at 45°), the aerodynamic effects of these would (one would think, at least) tend to couple together less efficiently, again producing a situation where the airflow is affected by the leading edge and trailing edge in sequential order, rather than all together.3
Does the rule that "effective sweep angle equals quarter-chord sweep angle" still hold for wings with very long chords, large differences between leading-edge and trailing-edge sweep angles, or both?  If not, how does one determine the effective sweep angles of these wings?

1: For a wing with straight-line,2 unkinked leading and trailing edges, calculating this angle is a matter of simple trigonometry.  For a wing made up of discrete segments with leading-/trailing-edge sweep angles varying from segment to segment (but remaining constant within each segment), the angle of the quarter-chord line should be calculated for each individual segment.  For a wing with curved leading and/or trailing edges, the quarter-chord angle is much harder to calculate by hand, but the overall rule still holds.
2: "Straight" here meaning "uncurved", not "unswept" or "heterosexual".
3: This would be accentuated by the fact that any wing of significant size with greatly-different leading-edge and trailing-edge sweep angles will, by necessity, have a very long chord along at least part of its span, due to the constraints of simple geometry.

Simplifications help to make valid approximations. But they should not be applied when the boundary conditions make those simplifications invalid. Your question is a good example.

This is because the aerodynamic effects of wing sweep are a product of the spanwise flow along the wing

Not really.
Take supersonic flow. Now the leading edge sweep is much more important than quarter chord sweep. A subsonic leading edge will cause leading edge thrust and thus lower drag. At transsonic speed most Mach effects are delayed because the acceleration and deceleration of air flowing over the wing depends on the contour along a line orthogonal to the local sweep angle. Yes, that means that highly tapered, swept wings will have more sweep effect near the leading edge and less towards the trailing edge.
In the end, the acceleration and deceleration of the local airflow depends on the local pressure gradients, that is, it is perpendicular to the local isobars. Those isobars follow the local sweep angle, at least at mid-span. At the center of the wing and at the wingtip the sweep angle of the isobars is lower than local sweep but good designs change the local airfoils such that the isobars are perpendicular to the local sweep angle. Of course, once this is done, the simplification with quarter chord sweep holds true over the full span.

Does the rule that "effective sweep angle equals quarter-chord sweep angle" still hold for wings with very long chords, large differences between leading-edge and trailing-edge sweep angles, or both?

No, because it is a simplification which helps to approximate the flow at mid-span and can be applied mostly to high aspect ratio wings. Once the wing becomes similar to a delta wing, it is no longer valid. On low aspect ratio wings the mid and tip effects will dominate the picture


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Why would control-surface balance horns cause problems at transonic speeds?
On many aircraft, control surfaces such as elevators, ailerons, rudders, etc., have "horns" which extend forwards of the surface's hingeline, wrapping around the tip of the wing or stabiliser, as seen on the left side of this image:

(Image via Rory McDonald here at AvSE, who won't say where they got it.)
This is beneficial since the control-surface area forwards of the hingeline cancels out some of the aerodynamic forces opposing control-surface deflection, and provides a place in which to put heavy things to mass-balance the control surface; both of these make the aircraft's flight controls more responsive and easier to use.
However, this comment by @PeterKämpf on this horny-control-surface answer suggests that balance horns would be poorly-suited for transonic aircraft:

Control surface balancing reduces hydraulic loads and, consequently, mass. It does make sense even with irreversible controls and is possible on the PC-24 because of its low maximum Mach speed. At airliner speeds the control horns will cause all kinds of trouble, but below Mach 0.74 they offer clear benefits.  [My emphasis; mach 0.74 is the beginning of the transonic range for most aircraft, and all modern jetliners cruise at transonic speeds.]

Intuitively, one would expect that, if anything, these forward-pointing balance horns would be a good thing in transonic flight; one of the problems that can arise when getting close to the speed of sound is the degradation of flight-control authority due to shockwave formation as high-speed air passes from the wing or stabiliser onto the control surface,1 and the balance horn, flying into clean air not disturbed by any fixed surface in front of it, would allow some control authority to be retained regardless of the effects of shockwaves on the remainder of the control surface.
Why would control-surface balance horns be a net minus at transonic speeds?

1: In practice, this is generally most evident as a reduction in elevator authority, which is a bad thing in most flight regimes, but especially so when flying near the speed of sound.


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Why are lower VHF frequencies used for towers?
So, why are usually lower VHF frequencies (around 118MHz to roughly 120MHz) used for towers and higher frequencies (around 130MHz to 136.975MHz) for ACC? What is the advantage of using lower frequencies for towers and higher for ACC since the wave propagates as a line of sight in both bands, that is, the maximum distance of the radio wave should be the same regardless if the frequency is 118MHz or 136MHz?

What are you basing your question on? I've never noticed such a pattern, have you analyzed  a database of frequencies?

I remember a few years ago one Croatian controller said that, I didn't think about it and it didn't bother me, but now I started my ATPL training and the same thing lecturer mentioned when teaching radiotelephony...I started looking at that more deeply and Croatian structure follows that perfectly. Now, I don't know if that's a coincidence or if that's just in our country..Zagreb Tower: 118.3 (backup 119.125), App: 120.7; FSS: 135.050; ACC: 135.8;
I mean there will sure be some exceptions,but it seems to me like there is some reason behind it and I don't know which,so I posted the question :/

In my opinion, I think there is something behind it, that lower frequencies are reserved for communication near a transmitting station (like towers, ground, etc.) unless all channels are occupied. If all channels are occupied than, of course, higher frequencies will be assigned to towers but I cannot back up that statement with some source or literature, so if someone has any information about it, I would appreciate it.

This may be an artifact of how your country allocates frequencies; it’s not true everywhere.

Ground control is very commonly 121.x

There's no specific reason why the lower frequencies are used for towers, but it's not a coincidence that they are grouped together (mostly, there are tower frequencies that are not at the lower end.)
Spectrum management is a complex effort to make maximum use of the fixed amount of frequencies available for use. In the US this falls under the FAA Spectrum Engineering Group AJW-1C.
One method of ensuring reliable safe communication is the intersite analysis. Basically, frequencies have to be used by multiple ground stations based on the airspace volume that the frequency needs to cover. This airspace is referred to as a frequency protected service volume (FPSV).
The goal is that an aircraft in the FPSV will only be able to communicate with the intended ground station. Since towers have a relatively small FPSV, typically 5,000' high (AGL) & 15 nmi radius, the tower frequencies can be reused closer together than a high altitude sector frequency that may cover a 150 mi radius at 45000 feet. A frequency that was used for both high altitude and a tower would be more difficult to de-conflict.
There are also other considerations that can affect the assignment of frequencies. It's all spelled out in the 6050.32B - Spectrum Management Regulations and Procedures Manual. It's rather lengthy at 475 pages. The details of VHF/UHF Comm frequencies is in Appendix 2.

+1 and ICAO Annex 10,  Volume  V,  §4.1.1 and Doc 9718 define actually a pre-allotment of the VHF band. In this pre-allotment, the aerodrome services are in the lower sub band, and the ACC in the higher sub band. Details are then refined at ITU region level.

in that frequency range (118-137MHz) you are correct, the percentage change in frequency from the bottom to the top of the range is small, so there will be no significant change in propagation characteristics, range, data rate, etc. going from 118 to 137. In addition, the same radio antenna can be effectively used across the whole frequency range, simply by trimming it for resonance in the center of the band.
So the real question is frequency allocation i.e., how it's decided who gets to operate where within a frequency band. Those allocations are often arbitrary (for example, there's no reason why one end of a ham radio band would be reserved for morse code as opposed to the other end of the band) but allocating different band segments for different uses reduces the chance of miscommunication.
For example, if tower/unicom was blended into ATC channel assignments, then a pilot could be one channel mistuned and find him- or herself talking to the tower when they were expecting to talk to an en-route controller.

May not have been an explicit consideration, but the bottom of the band is pretty close to Band II FM broadcast radio.
118MHz could interfere with 96.6MHz transmissions, the difference between them being 21,4 MHz, twice the usual 10.7MHz IF (intermediate frequency) This is known as the "image frequency". It's only problematic with a "high side" local oscillator (LO), 10.7MHz higher than the broadcast station. This problem could arise with any channel up to 108 + 21.4 = 129.4 MHz.
You don't really want interference sources flying over millions of people if you can help it. People living close to ATC transmitters may notice some interference, (which can be dealt with in other ways, e.g. choose a radio with a low side LO, which is immune to the problem) but they might have more pressing noise problems, like aircraft flying low overhead.
So, keep the potential interferers stationary, move the mobile transmitters above 129.5 MHz.
(source : I had a job designing a stereo FM receiver once)
Many years ago I used to hear LF navigation beacons (280 to 530kHz) on an AM radio, via the image frequency. One played "GLG" in very slow Morse code (presumably Glasgow) - these must have been switched off about 40 years ago.


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What is the fuel mass flow rate at stationary full power for the CFM56 5C2/3C1 engines?
Does anyone know what is the value of fuel mass flow rate when stationary at full power for CFM56-5C2 or CFM65-3C1 turbofan engines? I am doing a work on engine performance and it seems that only information I can't find is the fuel flow rate at conditions above. I will be very grateful to someone who can help me.

fuel consumption times thrust?

@user3528438 You mean TSFC? If you know the TSFC at conditions above I can calculate the fuel mass flow and would resolve the issue.

@ymb1 Pretty much. But -3C will also work as well I think. Any data about it would be helpfull.

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Why the gradient of $C_L$-$\alpha$ graph increases with the Mach number?
From the JAA ATPL Book 13 (Principles of flight) the gradient $\frac{dC_L}{d\alpha}$ of the $C_L$-$\alpha$ graph increases for higher Mach numbers. Can someone explain why this is the case?

This increase in lift curve slope with Mach is commonly predicted via the Prandtl-Glauert correction for 2D airfoil:
$$C_{l_\alpha}=\frac{C_{l_{\alpha_0}}}{\sqrt{1-M_\infty^2}}$$
where $C_{l_\alpha}$ is the 2D lift-curve slope, $C_{l_{\alpha_0}}$ is the 2D lift-curve slope in incompressible flow, and $M_\infty$ is the free-stream Mach number (Ref. Anderson, Fundamentals of Aerodynamics), and is accurate before local shock appears.
The physical reason is the magnification of pressure change due to density change (compressibility), as understood from the isentropic relation:
$$\frac{\rho}{\rho_\infty}=\left(\frac{p}{p_\infty}\right)^{1/\gamma}$$
where $\rho$ is density and $p$ is pressure (Ref. Drela, Flight Vehicle Aerodynamics).


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Is it possible to have the laminar flow above $Re=10^7$?
For pipe flow I came across the information that the onset of turbulent flow occurs at approximately $Re=10^3$ to $Re=10^4$, while for boundary layers on the airfoil the onset occurs in between $Re=10^5$ and $Re=5\cdot 10^6$. Therefore, since the onset of turbulent flow ranges greatly, is theoretically still possible in some cases to have the laminar flow above $Re=10^7$ on the airfoil, disregarding the laminar sublayer in every turbulent flow above the surface?
Any reference to books would be highly appreciated.

Only in a favorable pressure gradient, or when suction is applied to keep the boundary layer fresh.

The airfoil always starts with a laminar boundary layer; at some point on the airfoil surface, the boundary layer may transition to turbulent. In a uniform free-stream, two modes of transitions dominate:

Natural transition
Forced transition

a. Natural transition
Natural transition is the transition mechanism that we typically associate with Reynolds number. As flow moves along the airfoil, the local Reynolds number increases and any initial disturbances amplify. When the amplification increases past the maximum threshold that a laminar boundary layer can tolerate, the flow separates locally (laminar separation bubble) and reattaches as turbulent boundary layer.
The following graph (Ref. Drela, Flight Vehicle Aerodynamics) shows the growth of the local Reynolds number ($Re_\theta$) and the instability parameter ($\tilde{N}$) for a flat plate. Once $\tilde{N}$ increases past the critical value ($N_{crit}$), then laminar-turbulent transition occurs. Notice also how the point of transition is affected by the free-stream Reynolds number ($Re_\infty$): the higher the free-stream Reynolds number (which is what your OP is asking), the earlier the natural transition.

Image ref: Drela, Flight Vehicle Aerodynamics
Since the shape of the airfoil affects the pressure distribution, it also affects the boundary layer shape. You can modify the airfoil shape to delay the transition as much as possible in order to minimize skin friction drag. This is the design objective of Natural Laminar Flow (NLF) airfoils. But even on NLF-0215F at a $C_l$ of 0.65, for example, transition occurs at 61% chord on the upper surface, at a free-stream Reynolds number of 5 million.
b. Forced transition
Forced transition occurs when there is irregularity or protuberance on the airfoil surface. This can be due to manufacturing tolerances, contamination, or by design. Contamination can be due to rain, ice-buildup or even collision with insects.
Thus in practice, actual transition would occur much earlier than predicted via natural transition.

The Reynolds number is the ratio of a characteristic velocity and a characteristic dimension (never mind the viscosity for now). The key is the choice of the characteristic quantities. So, the onset of turbulence around Re=2300, what you often find in fluid dynamics textbooks, applies only to pipe flow, with the pipe diameter used as characteristic length scale.
The flow around airfoils is fundamentally different to the flow in a pipe.


text:
What is the difference between Vs and Vs1g?
As the name suggests, what is the difference between Vs1g and Vs?
If both of them are referring to the stall condition, why they aren't the same?
Which one is usually greater and why? Why do we need a distinction between them?

This is likely a V-n diagram problem...

This question would benefit from some context-- where is an example of the term "Vs1g" being used?

User Sidestick_n_Rudder posted on PPRuNe (emphasis mine):

Vs1g is the lowest speed, at which the airplane can maintain 1g, i.e.
level flight. It corresponds to the Cl max
Vs is the lowest speed attained during stall testing of the airplane.
The pilots were able to reach this speed lower than Vs1g, but the
plane was already losing altitude.
The way I understand it, the certification rules were changed at some
point, and new airplanes are required to base their performance on
Vs1g, whereas airplanes certified earlier were referencing to Vs.
However, this puts the new airplanes at a disadvantage, as Vs1g is
higher than Vs. To counter this, Airbus managed to convince the
certifying authorities, to allow for lower margins during performance
calculations (eg. Vref is 1.23*vs1g instead of 1.3), because the
airplane is speed protected anyway....
P.S. strangely enough, ATR-42/72 also use Vs1g and lower safety
margins, even though they are a conventional plane, not
speed-protected

I found another resource (AviationThrust) that mentions (emphasis mine):

Vs1g and Vs are two speeds majorly abbreviated for stalling
conditions.

Vs1g corresponds to maximum lift coefficient i.e. Clmax. Clmax is the condition when angle of attack is maximum and load factor is equal to
one.
Vs
corresponds to conventional stall condition. This is the time when lift suddenly collapses and load factor is less than one.

Vs is less than Vs1g

We usethe definition of lift from here:
$$ L = C_L \frac{1}{2} \rho V^2 S \tag{1}\label{eq1}$$
and the definition of the load factor:
$$ n = \frac{L}{W} \tag{2}\label{eq2}$$
Plugging \eqref{eq1} in \eqref{eq2} gives:
$$ n \cdot W = C_L \frac{1}{2} \rho V^2 S \tag{3}\label{eq3}$$
Rewriting for $V$ gives:
$$ V = \sqrt{\frac{n \cdot W}{C_L \frac{1}{2} \rho S}} \tag{4}\label{eq4}$$
Equation \eqref{eq4} can be satisfied in different ways:

Initially, there was only Vs, with no requirement on n, so set n to any value (with n<1 meaning falling from the air) to reach $V_s$

$$ V_{s} = \sqrt{\frac{(n<1) \cdot W}{C_{L} \frac{1}{2} \rho S}} $$

The regulations were updated to define straight flight, so set n to 1 and $C_L$ to $C_{L_{max}}$ to calculate $V_{s_{1g}}$

$$ V_{s_{1g}} = \sqrt{\frac{1 \cdot W}{C_{L_{max}} \frac{1}{2} \rho S}} $$
For $V_{s_{1g}}$, we always require 1-g condition, meaning that lift and weight are balanced. For $V_s$, there is no such requirement. By setting the requirement for 1-g, it is possible to more rigidly define the stall speed.
When slowing down in level flight, you will first experience $V_{s_{1g}}$, then the aircraft will start to fall (n<1) and you will reach $V_s$. Therefore, $V_s$ is lower than $V_{s_{1g}}$, see the image from this Quora link
(In the image the CAS arrow is in the wrong direction, right is lower CAS)

Nice overview! Informative...


text:
What is the difference between static and dynamic stall?
What are the static and dynamic stalls on the aircrafts wing? When would we encounter static and when dynamic one? What is the difference between them?

Static stall is what we typically think of, when we think of stall. Slowly increasing AOA and then loss of lift as the flow separates.
But pitch rate (rate of increase of AOA) has an impact too. At high pitch rates, the wing can go beyond normal stall AOA and still provide significant lift, but for just a short period of time. After that, the bottom falls out. This is dynamic stall.

Updated per request. OK?

yup, got my +1 now. Minor one, "google it" on SE often results in a catch 22 loop, as Google will likely show this question as one of the top results now, it likes stack exchange answers, so best to make answers as complete as possible.

These may be regional terms, I have not heard of static vs dynamic stall, I am a pilot in the USA. Airplane flight (heavier than air, lifted mainly by reaction of wings not engine thrust.) is inherently dynamic and so a stall must be as well.
I assume the question is actually asking about a typical low speed stall compared to an accelerated stall? An accelerated stall occurs when the  wing load is increased substantially above the airplane gross weight due to acceleration along the vertical axis(the airplane's vertical axis, not earth's); examples are pulling up sharply at cruise speed or while in a steep turn. Both types of stall happen at the same effective angle of attack but at different forward speeds.
Accelerated stalls can have a more sudden effect but generally recovery is also quick because the plane only needs to reduce the wing load; in a non-accelerated stall the forward speed is so slow that maximum lift is less than the weight of the airplane and some speed must be recovered before altitude can be maintained.

You are right, even a static stall is a bit dynamic. Maybe "quasi-static" would be a better word for it, but it is too long. Both the static and the accelerated stall are the same if entered with a slow speed decrease or load factor increase. At higher load factors, the asymmetry of the turn will make the accelerated stall differ more and more from the static stall. And a dynamic stall is yet another thing, as @MikeY correctly explains. This answer has more on this.


text:
Can Transition Altitude (TA) and Transition Level (TL) be equal?
In Europe, is it possible that Transition Altitude and Transition Level are equal, that is, TA=8000 ft and TL=FL080? In that scenario, at 8000 feet, what pressure (QNH or QNE) should be set in the altimeter? Is it possible that Transition Altitude is higher than Transition Level, e.g. TA=10,000 ft and TL= FL090? If the above scenarios are not possible, is there some minimum altitude of the transition layer that must always be assured?

Related: https://aviation.stackexchange.com/questions/60826/calculation-qne-and-qnh-with-flight-level-and-qnh-set-lower-higher-than-qne

The ATC gives a clearance of descend altitude 8,000 QNH 1011 then set the 1011 during the descent to 8,000.  If ATC says descend Flight Level 80 then keep 1013 in the altimeter setting and descend to 8,000.

Both situations you describe are theoretically possible. Saying that 8000 FT is "equal" to FL80 is a false assumption. Only in one very specific situation would that be true: when the QNH is 1013. Whenever the QNH is different than 1013, 8000 FT and FL80 would be two physically different levels.
In Europe, the TA and the TRL must always be separated by at least 1000 FT1. The distance between TA and TRL is called the Transition Layer. The common way to handle this is to have a fixed TA (often 5000 FT) and a TRL that varies, depending on the QNH, so that the Transition Layer is always between 1000 FT and 1500 FT "thick".
What this means is that on days with a low QNH, the TRL will have to be "higher" (numerically, not physically) and vice versa on days with a high QNH.
Counterintuitively, if the QNH is 1047 hPa or higher, 8000 FT and FL80 are separated by 1000 FT, which means that if the TA is 8000 FT, the TRL can be FL80. Two aircraft passing each other, one at 8000 FT (QNH 1047) and one at FL80 (QNE, 1013 hPa) would be separated by 1000 FT. It is, theoretically, possible to have a TRL that is "lower" than the TA (for example TA 8000 FT and TRL FL75) while the TRL is still physically 1000 FT above TA, but that would require an unrealistically high pressure.
Similarily, when the QNH is low, the TRL has to be "higher" to ensure separation. If the QNH is 1010, 5000 FT and FL60 are not separated by 1000 FT (the actual separation would be roughly 900 FT), so in that case the TRL would have to be FL65.
I work at an airport with a fixed TA of 3000 FT. The TRL is continuously updated based on the QNH, and typically lies in the range of FL35 at the lowest and FL50 at the highest. But, the actual distance between 3000 FT and TRL will always be between 1000 FT and 1500 FT.
1: ICAO Doc 7030 EUR

Do you maybe have any reference to a document that specifies the minimum height of the Transition Layer in Europe (at least 1000ft)?

@Darjan Page 31 (item 6.3.1.2) https://www.icao.int/EURNAT/EUR%20and%20NAT%20Documents/EUR%20Documents/EUR%20SUPPs%20working%20copy/Doc%207030%20(EN)%20-%20Edition%205,%20Amd%209+APP%20[EUR].pdf

@Darjan just to add to the above. Any differences or emphasis to this point will be found in each country’s AIP in ENR 1.7. For example the UK AIP ENR 1.7 - 3.4 states “In the UK, the first available Flight Level above the Transition Altitude is separated from the Transition Altitude by a minimum pressure interval corresponding to a nominal 1000 FT in the ISA” https://www.aurora.nats.co.uk/htmlAIP/Publications/2023-08-10-AIRAC/html/index-en-GB.html

The TA and TL will always be at least 1000ft apart to ensure that aircraft flying at each will always be properly separated.
The TA will be a fixed value to ensure terrain and obstacle clearance, while the TL will be the lowest reasonable value that provides this minimum vertical separation above the TA. However, “reasonable” includes not having any possible numerical ambiguity.
So, if your TA is 8000ft, then the lowest possible TL is FL085. Depending on the QNH, though, the TL may be higher.

Not true. TL could be 85 and still be separated from TA by 1000 ft, depending on the QNH

I work at an airport where we commonly have TRL FL35 (TA is 3000 FT).

@expeditedescent Fixed, forgot about VFR.

Nothing to do with VFR. And could you please add a source for your claim that with TA 8000 FT the lowest available TRL is FL85? I see no reason, with sufficiently high QNH, it couldn't be 80


text:
Why is Reverse thrust more effective at higher airspeeds?
It might seem logical at first but why is reverse thrust actually more effective at higher airspeeds?
If we assume that engine performance will not change during a 100 knot deceleration on landing roll-out, and engine is kept at constant reverse thrust throughout the whole landing then the engine will produce a constant force to slow the aircraft down. It won't be more effective at higher airspeeds because we are continuously redirecting a constant amount of air at specific velocity. Even if we include an engine performance change on landing roll-out, thrust will probably not change by more than a percent. Am I missing something?

FWIW, but perhaps not getting at what you're looking for, if you leave the engines at idle thrust, but just deploy the reverser doors (or whatever you choose to call them) that in itself will cause a significant at amount of drag over their closed position at high speed, especially if they're the old clam-shell doors. That drag will decrease as the airspeed across them decreases.

From "handling The Big Jets" by D.P.Davies.
Reverse thrust is much more effective at high aircraft speeds than low speeds for two reasons. Firstly, the net amount of reverse thrust increases with speed because the acceleration imposed on the (constant) mass flow is greater. This is because the aircraft forward speed is additional when in reverse thrust as opposed to subtractional when in forward thrust. Secondly the power produced is higher at higher speeds because of the increased rate of doing work. In this context it means that the kinetic energy of the aeroplane is being destroyed at a higher rate at higher airspeeds.

In simple terms when you deploy the reversers (clamshell or other) it is as if you were sticking both your hands out of a cruising car at 200 Km/h, obviously the slower you go the less effective the resistance on your hands becomes, all the way down to 0 Km/H were the resistance is nil. Same with the reversers....


text:
What is primary cause of shock induced separation (shock stall)?

In the video link above describing Prandtl Meyer Waves, lecturer stated (at 22:40) a very important fact on which he based his first half of the lecture. "Mach wave (mach lines) deflect the flow through an infinitely small angle".
Now, lets look at the flow over the airfoil. First the flow over the top surface will accelerate to supersonic speeds(and above) until it reaches the point(roughly) of maximum camber. Due to acceleration, mach lines will diverge and of course no shock wave will form. But, after the point of maximum camber, where adverse pressure gradient will be introduced and decelerate the airflow than mach lines will start to converge (coalesce). Once they do it the shock wave will form above an airfoil (picture below). If aircraft countinues to accelerate, the mach lines would intersect each other more lower and lower to the airfoil bringing a shock wave down, until it meets the airframe (this sentence is my assumption and I would like confirmation if it is correct). At that point(once it touches the airframe) the boundary layer will always separate and cause a shock stall(also my assumtion, will it happen always?).
My question is:
Is boundary layer separation caused by the extreme instantaneous pressure increase on the airfoil surface, or because of the convergence of mach lines and deflecting the airflow regardless of the pressure(and density) change? Or is it combination of both?
Also if the statement from the lecturer is correct, shouldn't the mach lines deflect the airflow sooner and not only at the point when they all merge together?

The simplest answer is that a shock wave for a boundary layer means a sharp adverse pressure gradient. You probably know already that adverse pressure gradient will thicken a boundary layer and thick boundary layers are more prone to separation. Also, if the adverse pressure grad. is so strong, it can immediately produce a detachment of the flow from the wall.

Add to this that a separation bubble forms in the near region and this interacts with the shock. All these phenomena are quite complicated and to explore their underlying physics Large-Eddy Simulations (LES) are used for the Navier-Stokes equations. This is a topic of Ph.D level research. If you want an introduction you can check this source and discover further on your own.


text:
Which popular turboprop engine feathers its propeller blades with each shutdown?
Chapter 12 of the FAA Airplane Flying Handbook, discussing multiengine flying, goes into quite a bit of detail in the propeller section about how most feathering-propeller-equipped engines on multiengine aircraft have a mechanism that prevents the propeller from feathering if the engine RPM is below a certain value, before mentioning that one particular widely-used turboprop model does not have this mechanism:

As just described, a loss of oil pressure from the propeller governor allows the counterweights, spring, and/or dome charge to drive the blades to feather. Logically then, the propeller blades should feather every time an engine is shut down as oil pressure falls to zero. Yet, this does not occur. Preventing this is a small pin in the pitch changing mechanism of the propeller hub that does not allow the propeller blades to feather once rpm drops below approximately 800. The pin senses a lack of centrifugal force from propeller rotation and falls into place, preventing the blades from feathering. Therefore, if a propeller is to be feathered, it must be done before engine rpm decays below approximately 800. On one popular model of turboprop engine, the propeller blades do, in fact, feather with each shutdown. This propeller is not equipped with such centrifugally-operated pins due to a unique engine design. [FAA Airplane Flying Handbook (FAA-H-8083-3B), Chapter 12 ("Transition to Multiengine Airplanes"), my emphasis.]

What is the identity of this standout, and what unique aspect of its design is the reason for this difference from other feathering-propeller-driving engines?

This is typical for the De Havilland Canada Dash-8, which is powered by Pratt & Whitney Canada PW100 engines. Other aircraft powered by that engine also seem to have their propellers feathered during shutdown. The propellers are only unfeathered after the engine has started (see e.g. this YouTube video of an ATR 72-600 starting its engines with the propeller still feathered).

From the Dash-8 Q400 manual:

Manual Feathering
Propeller manual feathering is used during engine shutdown by selecting the appropriate condition lever to the START & FEATHER and/or FUEL OFF.

(Dash-8 Q400 Manual - Propeller)
Normal shutdown procedure is to select the CL (condition levers) to START & FEATHER first, then let the engine run for about 30 seconds (with the propeller already feathered) and then select FUEL OFF:

POWER AND CONDITION LEVERS ..... DISC / START&FEATHER
[...]
COND. LEVERS (AFTER 30s) ..... FUEL OFF

(Dash-8 Q400 - Normal Procedures)

(Dash-8 Q400 Manual - Powerplant)
The manual does not state a reason for feathering the propeller on every shutdown. It does quickly reduce the propeller speed on shutdown and therefore allows the ground crew to approach the aircraft without much delay. Such a design is only possible on free turbine engines because otherwise high drag from the feathered propeller would require too much power during starting1.
The difference to other turboprops might be related to the design of the alternate feathering system:

Propeller Feather Pump
The Auxiliary Propeller Feathering Pump Unit provides an independent means of feathering the
propeller in the event of a failure of the primary means of feather. The auxiliary pump consists of
a 28VDC electrical motor driving an external gear pump which supplies a secondary source of
pressurized oil for feathering the propeller. The feather pump is also used for feathering and
unfeathers the propeller as required for maintenance purposes.

(Dash-8 Q400 Manual - Propeller)

1: For example, on the Dornier 228 (powered by single spool Honeywell TPE331 turboprop engines), the propellers are intentionally unfeathered for starting to reduce drag:

For ground starting, the propellers are held mechanically for minimum drag in a fine pitch position by
spring actuated start locks. [...] To engange them when the engine is not running the UNF-PUMP must be
activated with the POWER lever in REVERSE until the minimum propeller blade angle is reached.

(Dornier 288-100 POH - Airplane & System Descriptions - Propellers)


text:
Is it possible for a fighter jet to shoot itself down with an IR missile?
In this answer, it's mentioned that

However, IR missiles (AIM-9 Sidewinder, AA-11 Archer, MICA IR, ASRAAM) do not emit any EMR that indicates they're incoming; they use a passive FLIR sensor to identify and track the heat source they were told to kill (they don't even require a radar lock; the seeker can be cued to a pilot's helmet, or it can be "uncaged" and will lock on to the most significant heat source in front of it).

Emphasis added.
Given this, is it possible, given some very bad luck, for a fighter jet to fire an IR missile, only for it to get turned around somehow (for instance, by some sort of momentary combustion instability in its rocket motor, or by some very bad turbulence) and end up locking onto and shooting down the jet that fired it?

Reminds me of near the end of The Hunt for Red October, where a Soviet torpedo hits the sub that fired it. (Interesting, although certainly not authoritative.)

Seems like a more likely (but still involves highly questionable Hollywood-style physics) would be for the aircraft to drop the missile before the missile's engine fully engages.  Aircraft then moves ahead of missile, missile locks on to aircraft, boom.

Yes, but not in the way you're thinking
As a general rule, a missile shouldn't get turned around like that -- if it did so as a sustained maneuver, it'd be vastly short on energy and thus easy to evade in addition to being well out in front of the launching aircraft, making a "circular running missile" self-shootdown scenario improbable.
However, it's still possible to shoot yourself down with a missile, even if it never locks onto you, as Grumman test pilot Pete Purvis and his RIO "Tank" Sherman found out the hard way during testing of the F-14 Tomcat.  Stores separation can be an aerodynamically tricky business, and getting torched with rocket exhaust is a good way to find yourself on fire.
While an IR guided missile will have a smaller rocket motor than a big radar-guided missile like the AIM-7 that did Pete and Tank's test Tomcat in, it's not beyond the realm of plausibility that an IR-guided missile could still cripple an aircraft simply by failing to separate from said aircraft correctly during the launch sequence.

"While an IR guided missile will have a smaller rocket motor than a big radar-guided missile" tell that to the Russians who have missiles where the same missile body and motor are mated with different sensor packages.

@jwenting yeah, that is a point :)

Not really, it would take any of these missiles so long to turn 235 degrees (way more than 6-10 secs), which is generally the maximum burn time of most of their rocket engines. Plus, if it did somehow, get turned around 235 degrees, the energy (airspeed) bleed from such a large turn would leave it without any energy to intercept the launch aircraft even if by some miracle it acquired an IR lock

Actually happened during Sidewinder weapon separation test.   The missile launched the rocket motor stuttered, fell about 300 feet behind what I think was the F105 launching it and then the rocket motor ignited and went right up the engine exhaust.  No warhead so the pilot ejected successfully.  The event was captured by the chase plane and I saw the video in an Air Force Now film as an Air Force cadet in the early 1980s.

Nice one, I guess this story is less well known then the F-11 that shot itself down with its own gunfire.

Unlikely but possible
It might be possible assuming that the missile does not read or request any IFF signals. If the motor on a missile were to malfunction enough or the missile was to experience enough turbulence to turn it around then it would probably destroy the missile or leave it having a hard time tracking a target. Maybe if it was fired at an aircraft that the managed to get behind the firing aircraft this could happen.
There was a large number of similar events with the first homing torpedos, however sonar based homing has a larger angle of view that the sensor can pick up.

Highly unlikely. It is however easily possible for two aircraft armed with IR guided missiles, which can nowadays acquire targets in the head on aspect, to shoot each other down.


text:
Do any airlines have Braille emergency checklists?
In reading the report into the crash of UPS Flight 6, I noticed that one of the items of feedback from pilots in a simulated inflight-fire scenario utilising a smoke-filled simulator was (page 287 of the report):

In the thickest smoke, only Braille will help. What bigger font does to you, if there is any "break" in the smoke, it allows you to see more letters (group), therefore part of words. From that you can deduct what the action on the checklist is for the next step.

(Emphasis added.)
Which makes sense, as Braille, unlike most scripts, is read by touch rather than by sight, and is, thus, unaffected by limited or nonexistent cockpit visibility, while visual scripts, even with large, simple glyphs, have their legibility severely impaired by a smoky cockpit (in dense smoke, the pilot(s) essentially have to wait to do anything until the smoke momentarily lifts), and are completely useless in a zero-visibility environment.
With this in mind, do any airlines (passenger and/or cargo) have Braille checklists for use in situations of severely degraded cockpit visibility (such as a cockpit filled to the brim with dense black smoke)?

This seems like ridiculous folklore, nobody I know of that goes through private, commercial, or ATP training was ever required to learn to read braille. If you can't see a checklist 6" in front of your face, how are you going to see the instruments? I know for a fact there are no "braille instruments"...

I'm pretty sure "only Braille will help" is semi-sarcastic, meaning there is nothing useful that can be done to make checklists useful in the worst circumstances

A Braille checklist would require pilots to learn, practice, and be proficient at reading Braille in high-stress emergency situations. Does that seem like a reasonable burden to you?

I agree with the answer given, but wanted to add that most checklists for emergencies of this type have (bold faced, or asterisk) items that are required to be committed to memory.  Also, certain knobs, switches, and handles are shaped differently, allowing them to be identified and actuated by feel in a low or zero visibility situation.  (smoke, loss of cockpit lighting at night, etc.)
During simulator flight training in the Navy we used to be required to perform some procedures blindfolded, finding knobs by feel based on their relative position and shape.  Memorization and blindfold drills would be far more useful expenditure of time than learning braille to follow a written checklist.

It would be counter productive

It takes time to learn (especially touch reading)
It takes time to keep proficient

This is time spent on a skill useless for >99,999999% of flights
This is time NOT spent on more useful safety trainings
This wouldn't work anyway

It would be very slow to find the correct checklist then read it
If the aircraft is shaking, it would hinder your speed/prevent your from reading
You have to take your hands of from other controls to read

There is better alternatives
Simply having a recording/speech synthesis/speech command would solve it.

This is time NOT spent on more useful safety trainings

This one is a killer


text:
Why don’t aircraft have interlocks to prevent high-lift devices from being retracted when doing so would stall the airplane?
Most aircraft have high-lift devices on the wings (such as flaps, slats, droops, whatevs) to allow them to take off  and climb out at a reasonable speed instead of having to accelerate to near cruising speed while still on the runway; this enables the aircraft to use shorter runways, but also places it at risk of stalling if these devices are retracted too early in the climb (example).
Given the danger of retracting high-lift devices at too low a speed, why don’t aircraft have interlocks to prevent this from happening?

Most modern airliners have some form or other of mechanism that does exactly that, mostly for the slats (which are more important in stall prevention, anyway).
E.g. on modern Airbusses, there is a Slat Alpha Lock function which prevents slat retraction at high angles of attack, on the Boeings there is a Slat Auto-Gap function extending slats from mid to gapped position at high angles of attack.
Military aircraft often have automatic slat deployment to aid with high angle of attack manoeuvres, and even some vintage aircraft did have aerodynamically activated slat-like devices for stall prevention.
Any sensible reference suggestions welcome for any of these... cannot post mine.

Airbus, Airbusses, Airbusi? :)

@AEhere Surely it must be Airbi... but sacrificing readability for correctness is not always the best tack ;)

To add something actually useful, iirc automatic slats were a thing as early as the Mig-3 (late versions), maybe earlier. Sadly I don't have a source at hand right now

@AEhere Yes, they had them on the Messerschmitt Bf-109, too.

While flaps do slightly reduce stall speed, that is not their primary purpose.  Planes pass quickly through both Vs0 and Vs1 during takeoff or landing, and pilots are too busy to mess with flap settings during those few seconds anyway. The only reason we deliberately fly inside that range is to practice slow flight, stalls and spins, where such an interlock wouldn't be of any benefit.
During takeoff, low flaps may be used to increase lift and thereby shorten the takeoff distance on short runways or get into ground effect sooner on soft runways. In both cases, though, the climb afterward will be slower due to the added drag, so it's only done when necessary and then pulled out at best climb speed, i.e. well above stall speed.
During landing, high flaps are used to increase drag and thereby descend without gaining speed. You're not even close to stall speed until the flare, and raising flaps before then just means you'd land faster--not stall.
OTOH, while the benefit (if any) would be tiny, there is a very real risk that the device would fail somehow, such as preventing flaps from retracting at any speed, which is a serious safety problem. You would need a way to bypass it in such cases, and if you trust the pilot not to stupidly bypass it, then you might as well trust him/her not to stupidly need it in the first place.

"You're not even close to stall speed until the flare, and raising flaps before then just means you'd land faster--not stall." Not true for transport category aircraft. For example, the minimum safe airspeed for a 747 at landing flaps is more than 60 knots slower than the minimum safe airspeed with a clean no-flaps wing. If you were inside the outer marker of an ILS at flaps 25 or 30, the two typical flap settings for a 747-200, and at the proper speed for that flap setting and suddenly retracted the flaps, you would stall. Your first indication would be the stick shaker going off.

I don't know about the 747, but can tell you that the smaller business jets I've flown typically have the Vref set at 20% to 30% above stall speed.

@Terry I know the big iron has lots of calculations to determine V speeds based on weight, but the standard rule of thumb if no Vref is provided is Vref=1.3 * Vs0.  Vs1 is typically only 10-15% higher than Vs0, so unless you're below Vref, pulling the flaps shouldn't result in a stall.


text:
How is the control of airspace over disputed territories determined?
Reading this question about the division of airspace over the Persian Gulf reminded me of another potential issue regarding Persian Gulf airspace: the numerous territorial disputes in the area, such as the dispute between Iran and the United Arab Emirates over several islands (Abu Musa, the Greater and Lesser Tunbs, Sir Abu Nu'ayr, and Zirku) scattered over a large area of the lower end of the Gulf, and the unresolved issue of how far northwest the United Arab Emirates extend.
How is the ownership and control of the airspace over these and other disputed areas around the world (of which there are rather a lot) determined?

International air traffic control is divided into Flight Information Regions (FIR) by ICAO through international agreement. Normally this works fairly well. Even countries that are antagonistic to each other wish to maintain safe air traffic.
Currently there is one situation that causes some real consternation, though. That's the dispute between Greece and Turkey over northern Cyprus. ICAO officially assigns the region to Greek ATC, but the northern part of Cyprus, which aligns with Turkey feels that the airspace is theirs and operates a separate ATC that  is not in cooperation with the official ATC, causing confusion for pilots.
They often create long and ridiculous notams to state their point.

...Aaand it turns out it isn't just Cyprus that has these problems.  The same sort of situation occurs over the Crimea.

@Sean What a mess. Accident waiting to happen.


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How does ATC locate a stealth aircraft if its transponder fails?
Stealth aircraft are built to have near-zero radar visibility, so that, in a warzone, it'll be harder for the enemy to see it on their radars to shoot it down.
However, air traffic control also use radars.
Normally, this wouldn't be a problem, since the aircraft is carrying a transponder that continually broadcasts positional and movement data (so-called "secondary" radar returns), not just to ATC, but also to other nearby aircraft (so that they can avoid running into it).
If an aircraft's transponder fails, though, then ATC is reduced to tracking it via "primary" radar returns (literally bouncing radio waves off of it and seeing where they're reflecting from)... but stealth aircraft are specifically built so that they can't be picked up by most primary radars (at least without a lot of trouble).
So how does ATC - and, for that matter, all the other planes in the air who would rather not run into it - know the position of a stealth aircraft if its transponder fails?  Do they have the flight crew of the stealth aircraft continually read off map position, height, speed, direction, etc. data and try to use this to steer other planes away from it?

@ymb1: Presumably such an antenna could be deactivated/stowed when the stealth aircraft was in an active warzone...

ATC primary radar isn’t built with the idea of detecting military-type targets, so it probably just wouldn’t see a low-observable aircraft. They would have to rely on pilots to report their position.

It would be up to the military pilot to "see and avoid", ATC has no authority over military flights anyway, they usually call in/do routing as a courtesy to local ATC, but are not bound by it.

Theoretically, dopler radar can detect the aircraftsl's wake.

It's entirely possible and legal to fly a plane without working transponder. Here's a great video by Boldmethod. Their transponder failed when they were still on ground, it's ok for ATC. More communication is needed of course.

Military Stealth Aircraft are only stealthy when configured for stealth!  If a pilot wants to be seen (as in a failed radio or failed transponder), he only has to lower his landing gear, or open any bays, or deploy speed brakes, and his radar signature will immedately become much larger.

I think you are overthinking things a bit, since it doesn't really matter if it's a stealth aircraft or not.
Many places in the world do not have primary radar coverage at all, meaning that if a transponder fails - stealth aircraft or not - it will disappear from the radar screen. Controllers will then revert to providing procedural (non-radar) separation, based on position reports transmitted by the pilot.
Providing separation based on position reports is fairly normal, and procedures to use are established, so it's not a matter of "trying" to steer other planes away - it's a routine task.

Thanx, although in North America (where most stealth aircraft are likely to be flying in peacetime), at the very least, primary radar is, to the best of my knowledge, standard ATC equipment.

@Sean, primary radar is standard ATC equipment almost anywhere, but it is nearly useless when the secondary fails both because it is not able to determine altitude to any useful precision and because most of them have a lot of ground clutter that the interesting target will be quickly lost in.

Also in the USA military flights are usually controlled directly by ATC, but in other countries they may be under military controller who reserves some parts of the airspace for them and then no further coordination is needed.

@757toga: And ground radar would be completely impossible.

@757toga - That's assuming that MTI works as intended - that's not always the case. I've worked sectors that habitually picked up ground clutter from highways, watercraft, and other objects which should've been filtered out.
Primary can pick up a LOT of things controllers don't want. The major civilian (FAA) radar systems in the US have a toggle to turn off primary returns, and standard operating procedure is to work with it turned off unless needed to track a transponderless aircraft (or at least it was where I worked).

@757toga, my second-hand information matches that of Aluminum Showers—static objects are not an issue, but ground vehicles, windmills and similar still cause enough trouble that tracking primary returns is unreliable at best. Besides, not much use without altitude information anyway.

@JanHudec “Cessna 123AB, traffic over the train tracks, primary only, slow moving, could be a helicopter. [pause] Could be a train.”

It depends on the aircraft.  One example was on the F-117, which had removable radar reflectors on the sides of the aircraft, allowing the jet to be detected by primary ATC radars.

As to more modern stealth aircraft, it’s possible they use similar systems to ensure ATC tracking in peacetime.

Seriously? Them little things? Impressive.

A screw not properly seated on a maintenance panel can make the plane show up like a barn door on a radar.  Not the size, it’s the reflectivity that matters

These are called Luneburg reflectors, and are indeed still used by the F22 and F35.

It doesn't sound like those are deployable in flight should the transponder fail, unless I misunderstood.

@FreeMan:  They are normally installed when flying anywhere in ATC controlled airspace.  They are only removed for stealthy combat missions, or stealth training exercises.

When a B2 visited the UK for the Fairford air show a few years ago, it was always escorted by a pair of F15s, one on each wing-tip. So radar simply had to look for the 2 F15s, and the B2 was the hole in between them!

This does not provide an answer to the question. To critique or request clarification from an author, leave a comment below their post. - From Review

@CarloFelicione looks mixed, +3/-2 on votes and 2/3 reviews said looks OK. It certainly demonstrates a way to deal with the issue so I think it is valid and does provide an answer, though a reference would improve it considerably. That said, your answer has no reference either.

Actually, we can locate and track them the same way as if our radar fails. Get position reports, as needed, from the aircraft. Not as many trained non-radar controllers in the US as there used to be, though, and the best non-radar tracking device, the flight progress strip, has been supplanted by EDST.

As others have said, the most basic level of enroute air traffic control (i.e. not "near the airport" control) is what is called non-radar control in the USA and procedural control in the rest of the world. The idea in any case is to prevent collisions between aircraft using the air traffic control services, and this can be accomplished in two different ways:

If the aircraft can be depicted as targets on a radar scope, the controller issues instructions to keep the targets at least some minimum distance apart. If the targets don't get close on the scope it means the aircraft didn't get close in the air.
If the aircraft cannot be depicted on a radar scope, the controller issues instructions to ensure the aircraft are procedurally separated—they do not cross the same fix at the same time, or they are on different airways, or at different altitudes.

In the United States this is becoming very much a lost art, as radar sites and radar scope systems have become more and more reliable—certainly I have never had to do it. And operating without radar contact limits pilots in several ways, most notably the pilots cannot operate on "random" (point-to-point) RNAV routes—they must be on defined airways and report when they cross various points along the airway. Many instrument approach procedures are not usable unless the aircraft is in radar contact (indicated by "RADAR REQUIRED" in the approach plate), and of course other services like vectors-to-final and radar traffic advisories are impossible. But it is possible for ATC to ensure safe separation even if the aircraft is not shown on radar.
So to answer your direct question:

So how does ATC - and, for that matter, all the other planes in the air who would rather not run into it - know the position of a stealth aircraft if its transponder fails?

The pilots report their position to ATC, who update the flight progress strip with that information and use that to mentally track the aircraft. For more information see this answer about position report phraseology.


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Do any gliders carry GPWS?
Do any gliders come with a ground proximity warning system (GPWS)?  I realise that, for a glider, a GPWS probably wouldn't be able to help quite as much as in a powered aircraft, since a glider, by definition, doesn't have any engines which it can apply TOGA power on (unless, of course, it's a motorglider), but it could still warn the pilot if they were descending too rapidly for their altitude or were about to land with the glider in an unsafe configuration, thus allowing the pilot to take the necessary actions for avoiding a crash (for instance, by pulling back on the yoke, jettisoning ballast, lowering the flaps and landing gear if the glider in question is so equipped, etc.).

Seems like this would be more of a distraction than a benefit.  As my first flight instructor was fond of saying, "Get your head out of the cockpit!"

Many glider pilots have a GPS flight computer that usually has terrain loaded in its database and can alert a glider pilot of terrain.  This is not quite a GPWS but is cheaper and uses a lot less power (glider pilots rely on batteries for their electrical system). Most GA aircraft don't have a GPWS either and usually rely on the GPS solution.  ForeFlight has a nice feature called Synthetic Vision which can allow you to see terrain in IMC conditions and is much cheaper than a GPWS.

GPWS is mostly useful for aircraft flying in instrument conditions or at night when there is low visibility preventing a good view of the ground. Gliders aren't designed to fly in these conditions and they don't have the required instruments. They should be able to see the ground and avoid it without GPWS.

Plus, don't gliders usually have not much of an electrical system?  I would imagine that radar is kind of power hungry.

Note that the second answer in your link shows that you can, in some jurisdictions at least, fly a glider in clouds.

I found this video of a glider equipped with GPWS, however, I doubt that such a system is very useful in glider, that isn't very complex and would consider it a gimmick.
Also because the sounds in the video are Boeing GPWS sounds, so it's probably self-built for fun.
So, to summarize: There surely are some self-built systems for fun and coolness, but I don't think there would be enough demand and need for a commercial system.

A glider will be usually flown according to VFR. So it is basically the task of the pilot to maintain adeqaute distance to terrain. The distance can be rather low, flying in mountains. As already mentioned, current onboard flight computers include terrain databases, that will provide you with AGL. Software like XCSoar will perform a final glide calculation taking terrain into account, telling the pilot if glide path will be in conflict with terrain.
More important than GPWS is obstacle collision avoidance, as obstacles like power lines or cablecars in mountains are hard to see. The popular traffic awareness system Flarm can optionally host a obstacle database and then will provide apropriate warnings.


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What is subcritical and supercritical boundary layer?
What is subcritical and supercritical boundary layer?
What is the difference between them?


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Is high speed buffet possible at supersonic speed?
Since high speed buffet is caused by boundary layer separation due to shock wave and at supersonic flight there is no shock wave directly onto the airfoil (it is attached to the trailing edge), is high speed buffet then possible?

All pilots who transited into supersonic speed on older designs which were not perfected for transsonic flight reported that suddenly all that shaking and vibrating stopped and flight became smooth again as soon as they went over Mach 1.
So the answer is no. With supersonic speed the buffeting stops.


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Hydromechanical/Electronic fuel control system
Can someone explain the picture below which represents hydromechanical/electronic fuel control? What's the function of the solenoid on the picture below? What are the differences between solely hydromechanical and hydromechanical/electronic fuel control? What exactly does the "electronic" part of the hydromechanical/electronic fuel control do? What engines had hydromechanical/electronic fuel control and which ones had hydromechanical?

Picture is from EASA Modulo 14 Propulsion (B2) but it does not provide any useful explanation.

This is kind of broad.  Can you reduce it to just a single question?

Posting a picture with 20+ labeled items & asking for "can you explain this" is a rather broad, open-ended question. Someone highly knowledgeable about the subject could probably write pages of explanation about it. Please formulate a more specific, focused question.

I can give you a generalized answer and most of the details you are looking for are hopefully buried here.  You can think of it as 3 broad categories or generations of engine fuel control.  First was pure hyrdomechanical, which used mechanical devices to regulate power setting and acceleration.  The very first systems like on the Junkers Jumo and other early engines was not much more than a needle valve connected to the throttle.  The pilot's eyes and brain using data from the gauges was all that prevented stalling/surging, overtemps, etc, needing very careful throttle movement and monitoring.
The first modern hydromechanical fuel control systems in the early 50s used mechanical regulating and governing devices so that when the throttle was moved it more or less commanded a power setting and the regulating devices would manage the amount of fuel admitted.  This allowed things like "slam acceleration" where the engine could take sudden throttle changes smoothly, which was a no-no on the very early jet engines.
In the 70s, the idea of adding electronic control to fine tune fuel management got started and you could call this the second generation.  This takes the traditional hydromechanical fuel controller and fine tunes the final fuel schedule in a way that can't be done mechanically.  This is done by software (an Electronic Control Unit or ECU) that drives a metering device through a torque motor (a torque motor is like a rotary actuator that moves in proportion to a voltage applied and springs back to zero if the voltage is removed).  There is still the old fashioned hydromech fuel controller and a baseline fuel schedule, but this is bypassed by an overlaid electronic fuel schedule, controlled by the ECU through its torque motor.
Most non-FADEC engines built after the 70s have this electronic fuel control, better thought of as electronically supplemented fuel control.  When the electronics are turned off in this type of system, the engine reverts to its baseline hydromech fuel schedule.
The last is FADEC where most of the mechanical regulating devices are discarded and the entire management is by software controlling the entire metering system.
In the diagram what you seeing is a hydromech system with electronic control typical of an engine in the 80s.   The electronic control is via the "torque engine" (or torque motor) on the far right.  Everything else is standard hydromech controls.  The Manual Mode Solenoid controls which has authority over the fuel scheduling; the hydromechanical components via the "mechanical mode cam" for manual mode, or the "metering head sensor" controlled by the torque motor for electronic mode.
In the cockpit if you turn the ECU off, the Manual Mode Solenoid transfers fuel control to the Manual Mode Cam which moves with the thrust lever.  In manual mode the engine will roll back somewhat, and to get the same thrust setting you have to move the thrust lever up a bit.  Also the engine will be a bit more sluggish in responses because it's basically reverted back to an older generation non-ECU engine.
Early non-FADEC models of the GE CF-34 turbofan and the PW-120 turboprops use this type of fuel control (and I assume most engines of that vintage).
You could think these ECU controlled engines as a bit like early fly-by-wire control, where you had a manual reversion mode with the traditional control linkages there in standby.  A FADEC engine is like full authority FWB with no mechanical controls other than those components operated directly by the computer.


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How can a single misplaced screw so massively increase a stealth aircraft's radar cross-section?
In this comment on one of the answers to an earlier question of mine, @CarloFelicione points out how even the tiniest of imperfections can make a stealth aircraft much more visible to radar:

A screw not properly seated on a maintenance panel can make the plane show up like a barn door on a radar. Not the size, it’s the reflectivity that matters

How is it possible for a single loose screw to increase a stealth aircraft's radar cross-section that dramatically?  Why does it show up like a barn door, rather than like, well, a screw?

It's not the screw but the edge of the panel which causes the RCS increase.

I'm pretty sure the original source for this piece of trivia is "Skunk Works: A Personal Memoir of My Years of Lockheed" by Ben R. Rich. A worthy read if you're interested in aviation history; light on the technical side but enlightening on the management behind Have Blue. The screw comment is about a single test where the prototype had a weirdly large RCS due to some screws not being seated properly.

In RCS, size does not matter once you move beyond the scale of the wavelength of the radar. Therefore, the RCS of a panel with an edge length of 10 cm is (theoretically at least) the same as one of a panel with 1 m edges. The screw must be so poorly placed that it will cause a larger panel to stick out so it has no electrical contact, with the edge pointing in the right direction, then this massive increase in RCS is indeed possible.

It is the electrical discontinuity that does the damage, two metal sheets in firm contact look like one surface, two sheets with a slot because they are NOT in firm contact looks like an RF antenna at a frequency where the slot length is a multiple of a 1/4 wavelength. Search term "Slot radiator".

Hmm, why does this work differently from visible wavelengths? If I shine light on an object and measure how much I get back, larger objects will generally reflect more light, a relation that continues far beyond the wavelength of my light. Are microwaves different somehow?

@HenningMakholm I think the point is that much of low radar cross-section is about reflecting the incoming energy away from its source. If part of your plane is acting as an antenna, it's radiating in all directions

@HenningMakholm That isn't really true when you start dealing with objects 100 nanometers across or smaller, where they don't really "reflect" light so much as "scatter" it. The physical characteristics of such small particles can make a large difference in how they scatter light.

@PhotoScientist: The claim in the answer is that for scales larger than the wavelength of the radar, bigger objects do not reflect more of the radiation. I have no trouble believing that small object give an attenuated return; I have trouble believing that large objects would not reflect more the larger they are.

@HenningMakholm You're right. This answer I think confuses beam width with wavelength. See figs 2 and 3 here. The loss of continuity caused by a panel gap is a near-field effect for which there is no direct analogy in optics/quasi-optics, so you cannot use general RCS or diffraction concepts to explain it.


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Used commercial aircraft engines
We are a scrap company and have experience with dismantling and scraping old commercial airplanes. Now we got an offer to receive an aircraft for its last landing, with engines, etc...I assume that engines are not possible to re-sell and they just want everything to be scrapped? any tips?

Who is "they"? Do "they" owned the engines, or had leased them to the engine manufacturer?

What type of aircraft are we talking about?

When you say you are a scrap company does that mean you are just really interested in harvesting the metal, etc. rather than recovering avionics, etc.? Reselling aviation parts is an involved process since all things aviation are tightly regulated. There are companies that operate solely as "aviation junkyards" that resell used parts. If the owner is contacting a scrap company they probably believe there are no useful parts, otherwise they would probably contact a company like Dodson that buys aircraft to resell parts.

we are waiting for more info on the type of the aircraft, but they are from 80 to 150 seats aircraft.

I dont know why they are deleting my comments/questions. I thought this forum is broad minded, but than I see that my question is marked as "not aviation related", than what? I have to explain certain things before asking a question, and when I ask a question, people ask me about more info, and I post more info in a manner of comment and a question. If you people are so strict, than I will go somewhere else and discuss without pressure.

They will have some value if you also receive the aircraft maintenance logs.  Components are virtually useless without a way to trace age, time, past service/repairs, and compliance with any service bulletins (SB) or airworthiness directices (AD).

we dont know at the moment, we just got in inquiry, mostly they are interested in licencing and documentation for dismantling for which we have.